In general, many current spacecrafts operate with zero momentum bias and use a stared thermostatic earth sensor assembly (STESA) or beacon and a gyro for on-station attitude determination. The earth sensor provides absolute roll and pitch attitude measurements for attitude update and gyro calibration, and the gyro provides relative roll, pitch and yaw attitude information. A sun sensor assembly (SSA) is utilized for providing absolute yaw measurement for attitude update and calibration of the gyro.
Each orbit period of a spacecraft orbit is operationally subdivided into two time durations: one is referred to as “gyro calibration” and the other is referred to as “gyro compassing”. During gyro calibration, 3-axis attitude measurements are used to calibrate the gyro and update the spacecraft attitude. The measurement data collected from the earth sensor and the SSA provide a 3-axis attitude update, which is also used to calibrate the gyro bias, and other gyro parameters. Gyro calibration occurs during approximately 4 hours of the day, which is when the sun is in the field-of-view (FOV) of the SSA and a good geometric dilution factor exists for accurate attitude determination by the SSA.
During gyro compassing, which generally occurs for approximately the remainder 20 hours of the day, 2-axis attitude measurements are used to estimate the spacecraft attitude. The earth sensor provides a 2-axis attitude update that includes roll and pitch measurements. The roll gyro sensor is used to estimate the yaw attitude by gyro compassing. Yaw attitude estimation is generally not precise and has associated yaw attitude error that increases over time due to roll gyro drift.
The yaw attitude error is often accounted for and corrected when the sun comes into FOV of the SSA. The resulting total yaw error accumulated during gyro compassing, which is the difference between the yaw attitude measured by the sun sensor and the yaw attitude determined by gyro compassing, is detected and a large estimated yaw attitude error signal is generated in response thereto. The duration of the large yaw attitude error signal is referred to as the yaw transient period. This yaw error accumulated during gyro compassing is erroneously used to calibrate the yaw gyro bias and causes undesirable transient in yaw gyro bias estimation. The erroneous yaw gyro bias also propagates over time into an undesirable yaw error until the yaw gyro bias converges back to its correct yaw gyro bias. Through attitude estimation and gyro calibration procedures the yaw attitude is corrected over a considerable amount of time.
Also, the yaw axis of the spacecraft is orbit-rate-coupled to the roll axis such that the large yaw error causes a significant roll error during the yaw transient period. The roll error causes the roll gyro bias to be calibrated incorrectly. As with the yaw attitude, it also requires a considerable amount of time to converge back to the correct roll gyro bias value.
In addition, given the limited time of sun in the FOV of the SSA, the roll gyro bias may not fully converge to the optimal roll gyro bias value. In other words, a small amount of error remains in the corrected roll gyro bias when the spacecraft switches to the gyro compassing. This is especially true when the period that the sun is in the FOV is not long enough with respective to the gyro calibration time constant. Everyday this undesirable yaw error transient is repeated, which persistently degrades the spacecraft pointing in gyro compassing.
Furthermore, it is inherently difficult to provide a clear FOV of 120° by 120° for a wide FOV SSA, due to the intrusion or blockage from other spacecraft components, such as antenna reflectors, thermal radiators, and solar wings, especially tilted solar wings. This further reduces the amount of usable time that is available, when the sun is in the FOV, for gyro calibration. Notwithstanding, a desired quality level of the geometric dilution factor needs to be maintained for this reduced amount of time. Thus, the available gyro calibration time is reduced and quality of the gyro calibration is degraded.
Moreover, gyro calibration is currently performed at one thermal condition corresponding to the thermal distortion between the earth sensor and the sun sensor for a small section of the orbit of the spacecraft, while gyro compassing is performed under other thermal conditions for the remainder of the orbit. Roll gyro bias is calibrated to account for the diurnal thermal deformation of when the sun is in the FOV of the SSA. The diurnal thermal deformation negatively affects yaw pointing of the spacecraft during gyro compassing. As a result, the calibrated roll gyro bias never satisfies the thermal conditions experienced during gyro compassing and the yaw error never converges to zero.
What is more, the spacecraft may operate in an SSA failure mode that can cause the spacecraft to loose fine mode data in turn causing the spacecraft to operate in a coarse mode. The coarse mode can be insufficiently accurate for attitude determination and gyro calibration.
Thus, there exists a need for an improved satellite attitude estimation and gyro calibration system that minimizes yaw attitude estimation error during gyro compassing, that utilizes a minimal amount of calibration time, that accounts for varying thermal conditions experienced during orbit of a spacecraft, and that can utilize SSA measurements in a coarse mode.